The compressor of an aircraft turbine engine typically includes arrays of rotating blades and stationary vanes or radial impellers arranged in stages along an axis of rotation. The form and angular arrangement of these airfoil shapes are such that rotation of the compressor rotor causes incremental compression and movement of air longitudinally through the compressor and into a pressure vessel called a combustor. A scheduled amount of fuel is metered into the combustor, and is burned with the compressed air to yield an energy transfer in the form of a high velocity gas flow. This flow is directed to a multistage turbine assembly that drives both the compressor and an engine output function which produces the required power or thrust.
Compressors of aircraft turbine engines are known to be subject to unstable operating conditions referred to as "stall" or "surge". Surge is an industry wide problem which is manifested by aerodynamic discontinuities within the flow path of the compressor, such that one or more stages of the compressor pump little or no air. The precise causes of surge are not known. However, it is believed that surge can be induced after individual or groups of stages have experienced stall conditions that result in a flow separation on the airfoil surfaces. This flow discontinuity may be caused by an inlet distortion or by a particular energy transfer in the combustor which creates a pressure and flow reversal through the compressor.
Surge generally is a transient condition which will clear itself after a short period of time. In many instances, surge manifests itself in the form of one or two "popping" noises and a minor vibration. This type of surge can typically occur during a slow steady state deceleration corresponding to a fuel control power level movement of 1.5.degree. to 5.0.degree. per second which establishes rate-of-change of compressor speed. For example, this surge may occur as the pilot retards the power lever and begins his descent into an airport. Although power lever movements which are slower than 1.5.degree. per second or faster than 5.0.degree. per second generally avoid this type of surge, it is extremely difficult for the pilot to know exactly how fast the power lever is being moved. Attempts have been made to incorporate a friction damper into the power lever to control its velocity. However, aircraft manufacturers feel this apparatus is too cumbersome, and pilots are reluctant to accept such a constraint on their control of engine power. Although the audible aspects of this low energy surge may be annoying, there is generally no structural harm to the engine.
Certain high energy surges can last longer and result in a noticeable jolt, a loud "banging" noise and possibly a continuous increase of gas temperature in the multistage turbine assembly. The forces accompanying this latter type of surge can affect the structural integrity of the compressor over an extended period of time. In certain instances, the pressure discontinuities in the compressor can result in an immediate nonconcentric operation of a compressor, such that the tips of the rotor blades may be urged into contact with the stationary vanes in the compressor. Hence, immediate and substantial damage to the compressor is possible. This type of surge typically occurs during small but rapid changes of power lever movement when minor adjustments are made to the speed of one or more engines in the bleed closed operating range.
Most turbine aircraft engines include bleed valves which avoid surge by releasing air from one or more stages of the compressor during certain operating conditions. In this manner, an acceptable flow and pressure distribution can be maintained across the various stages of the compressor. The specific ranges at which bleed takes place typically are related to the predefined regions of the engine fuel schedule.
One apparatus for bleeding the compressor to offset surge is shown in U.S. Pat. No. 3,006,145 which issued to Sovey. U.S. Pat. No. 3,006,145 is directed to a complicated mechanical arrangement where air can be gradually bled from the compressor in response to the corrected rotor speed, rotor acceleration and to some indication of fuel rate transients, such as the movement of the power control lever.
Another complex apparatus for controlling compressor surge is shown in U.S. Pat. No. 3,971,208 which issued to Schwent. The disclosure of U.S. Pat. No. 3,971,208 is directed to a complex electromechanical fuel control system which includes a surge valve to gradually bleed a controlled amount of air from the low pressure compressor in response to the surge schedule, the power lever position and a feed back of the fuel flow signal to the fuel flow regulator.
Other techniques for dealing with compressor surge are shown in U.S. Pat. No. 3,103,785 which issued to Williams et al and U.S. Pat. No. 4,164,033 which issued to Glennon et al.
Typically the specific electrical or mechanical apparatus for operating the engine and controlling compressor surge is designed in response to bench testing of the particular turbine engine at seal level conditions. However, it is known that surge becomes a more frequent and severe problem at higher altitudes due to lower air density and increased viscous drag (Reynold's number effects) and the tendency of boundary layers to effectively change the aerodynamic characteristics of the compressor. Specifically, although the normal air bleed schedule will avoid surge during a gradual deceleration of an engine at sea level, the slow steady state deceleration of a turbine engine at altitude typically may cause an audibly noticeable low energy surge condition. As another example, small but rapid changes in fuel flow are much more likely to cause severe and potentially damaging surges at altitude than at sea level conditions. In view of the increased frequency and severity of surges at high altitudes, it has been a problem to design turbine engines which operate rapidly and efficiently at low altitudes and which avoid surges at high altitudes.
Accordingly, it is an object of the subject invention to substantially eliminate compressor instability referred to as surge.
It is another object of the subject invention to specifically eliminate compressor surge which occurs at altitude.
It is a further object of the subject invention to eliminate the compressor surge which tends to occur during a slow steady state deceleration at altitude.
It is an additional object of the subject invention to eliminate the compressor surge which is attributable to small, abrupt changes in the fuel flow caused by rapid changes to the power lever position at altitude.
It is still another object of the subject invention to provide an apparatus for reliably eliminating compressor surges at altitude without substantial modifications to the aircraft turbine engine or existing control system.